Spacecraft payload ejection system

ABSTRACT

There is disclosed a Payload Ejection System (PES) able to store any set of payloads for launch and eject that set of payloads at a controlled speed with a low tumble rate while accommodating any offset centre of mass within a restricted volume. The need for ballasting or balancing is eliminated thus freeing up the design space for these payloads. Insensitivity to centre of mass location is enabled by the use of a deployment hinge assembly arrangement which uses two or more non-parallel folding hinge arrangements that allow for linear motion of the output link in one direction while restricting the motion all other directions. One embodiment of the current PES concept uses four (4) hinge panel assemblies, selected to provide optimal stiffness around the entire mechanism. The stiffness of the PES is integral to managing offset centre-of-mass locations by allowing the mechanism to translate the effective force vector to the center of mass location.

FIELD

The present disclosure relates to systems for hosting payloads on a hostspacecraft for the purpose of carrying the payload to orbit and ejectingit from the host spacecraft in a controlled manner. The payload ejectionsystem is the entire system which enables the ejection of a payload froma host spacecraft in a controlled manner. Ejection in a controlledmanner is defined as an ejection whereby the ejected payload has noangular momentum or linear momentum transverse to the ejection axis atthe time of release.

BACKGROUND

When satellites are launched to orbit (regardless of orbit type) thereis often some launch vehicle mass and volume capacity that is not used.One purpose of the system disclosed herein is to use this surplus volumeand mass capacity to deliver additional and separate payloads to orbit,from where the payload can proceed with its intended mission. Thisconcept of delivering hosted payloads to particular orbits is describedin “DARPA Phoenix Payload Orbital Delivery (POD) System: “FedEx to GEO”,Dr. Brook Sullivan et al, AAAIAA Space 2013 Conference and Exposition,Sep. 10-12, 2013, San Diego, Calif. As described in this paper, apayload includes but is not limited to such space systems as anothersmall (micro or nano) spacecraft, replacement materials (e.g. fuel) toreplenish another satellite, replacement components for on-orbitservicing repair of another spacecraft, components for in-space assemblyof a new space system or spacecraft.

Current orbital payload ejection systems require that the payload centreof mass be closely aligned with the centre of force of the ejectionmechanism or else significant tumble rates (undesired angular rates andtranslational velocities transverse to the ejection axis at the time ofrelease) are created at ejection, which is almost always considered avery negative condition. Accommodating an offset between the mechanismcentre of force and the payload centre of mass that remains unknown, butwithin a prescribed volume, at launch allows for increased flexibilityin accommodating payloads. This flexibility is particularly beneficialif there are multiple payload parts that may have specific packagingrequirements or irregular shapes. Similarly, endeavouring to make theprescribed volume for the centre of gravity as large as possiblemaximises the payload accommodation flexibility.

The current state of the art either uses an array of separation springs(e.g. the commercially available Lightband™) that can induce asignificant tumble rate if the center of mass is spaced from theejection mechanism geometric center, or a guide rail system (i.e.Pico-Satellite Orbital Deployer PPOD) for very small payloads(nano-sats) that does not scale well to larger payloads—in excess oftens of kilograms up to a few hundred kilograms—and, further, would beat risk of jamming or binding upon release.

Existing ejection methods are unable to eject a payload with an offsetcenter of mass without causing the payload to tumble. This is a resultof the ejection technique; many existing methods exert a force or forcesthat are on, or average to, the geometric center of the ejection device.If the center of mass of the payload is offset from this geometriccenter of the ejection device, the payload will tumble. A commontechnique in the industry is to use springs to eject a payload. If thepayload center of mass is offset from the geometric center, the forceupon the springs is not evenly distributed. This results in the payloadtumbling when the springs are released.

SUMMARY

The present disclosure provides a system and method of ejecting apayload from a host spacecraft in a microgravity environment. The systemand method does not require the payload have a geometrically centralizedcenter of mass. It also minimizes the tumble rate of the ejected payloadwhile being insensitive to the location of the centre of mass of thatpayload.

There is disclosed herein a method for controllably ejecting a payloadfrom a host spacecraft, comprising:

attaching a payload to a payload assembly, releasably affixing thepayload assembly to a payload release plate of a payload ejectionmechanism, said payload ejection mechanism including a releasemechanism;

said payload ejection mechanism including at least two deploymentassemblies coupled to said base plate via a first hinge coupling at oneend thereof to the base and at another end thereof to said payloadrelease plate via a second hinge coupling, said at least two deploymentassemblies being extendable between a stowed position and a fullyextended position, said first and second hinge coupling each having ahinge axis, said hinge axes being non parallel to each other so thatupon deployment said at least two deployment assemblies are constrainedto deploy such that a plane of said payload assembly remains parallel tosaid base plate;

wherein said payload ejection mechanism accommodates payloads havingcenters of mass not coincident with the geometric center of mass of saidpayload release plate in order to eject said payload assembly away fromsaid host spacecraft such that an effective force vector generated bysaid payload ejection mechanism acts through a center of mass of saidpayload at a moment of release of said payload assembly from saidpayload ejection mechanism regardless of the location of said payloadcenter of mass with respect to the geometric center of said payloadrelease plate; and

ejecting the payload assembly by activating the release mechanism todeploy said at least two deployment assemblies from their stowedpositions to their fully extended positions to force the payloadassembly away from the host satellite.

The at least two deployment assemblies are at least two deployment hingeassemblies, said at least two deployment hinge assemblies each includingtwo hinge plates hinged together along a common mid-axis axis along thelengths of each of said hinge plates, and one hinge plate of each ofsaid at least two deployment hinge assemblies being hinged to said baseplate along said first hinge coupling and defining a lower hinge axisand the other hinge plate being hinged to said payload release platealong said first hinge coupling and defining an upper hinge axis, andwherein each said common mid-axis, lower hinge axis and upper hinge axisof a given deployment hinge assembly is coplanar with each respectivesaid common mid-axis, lower hinge axis and upper hinge axis of all otherdeployment hinge assemblies.

The payload ejection mechanism may include at least four deploymenthinge assemblies with a first of two pairs of said at least fourdeployment hinge assemblies being in opposed relationship, and with asecond of two pairs of said at least four deployment hinge assembliesalso being in opposed relationship such said at least four deploymenthinge assemblies form a rectangular shape.

The at least one launch lock mechanism is for restraining and securingsaid payload ejection system and said payload assembly until a commandedtime of deployment at which time the launch lock mechanism isdisengaged.

The payload ejection mechanism includes an actuator to actuate saidrelease mechanism.

The actuator may include springs to actuate said payload ejectionmechanism.

The actuator may include a motor to actuate said payload ejectionmechanism.

The payload assembly may be permanently attachable to a payload.

The payload assembly may be releasably attachable to a payload.

A further understanding of the functional and advantageous aspects ofthe disclosure can be realized by reference to the following detaileddescription and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described, by way of example only, withreference to the drawings, in which:

FIG. 1 is an underneath oblique view of the payload ejection system(PES) 100 in the stowed configuration. It shows the relative positionsof the payload assembly 200, mechanical mounting assembly 400, with thebase plate 430 and the attached launch lock assemblies 410 and releasemechanism 460.

FIG. 2 is an end view, taken along arrow 2 in FIG. 1, of the payloadejection system 100 in the stowed configuration illustrating where thesections for subsequent figures are taken.

FIG. 3 is an isometric view of the payload ejection system 100 in thedeployed configuration. It shows the relative positions of the payloadassembly 200 and the payload ejection mechanism (PEM) 300 whichcomprises the mechanical mounting assembly 400, the deployment hingeassemblies 500 and the payload release plate assembly 600.

FIG. 4 is an end view of the payload ejection system 100 along the arrow4 in FIG. 3, in the deployed configuration illustrating where thesections for subsequent figures are taken.

FIG. 5 shows an isometric view of possible locations for the payloadejection system 100 to reside on the hosting spacecraft 700, such as anunused battery bay 720 or an unused portion of an outer surface 710.

FIGS. 6a and 6b illustrate the rotation or tumble 770 caused by alateral displacement 790 of the payload 800 centre of mass 750 from thecentre of an applied ejection force 760 or from the geometric centre 785of a plurality of ejection springs 765.

FIG. 7 shows an isometric view of a simplified linkage used in thepresent system using two (2) pairs of hinge plates, with simplifieddepictions of the upper hinge plate 501, lower hinge plate 502, upperhinge pin 503, mid hinge pin 504, lower hinge plate 505, base plate 430and payload release plate 610.

FIG. 8 is a section along the line 8-8 in FIG. 2 illustrating therelative positions of the major assemblies in the stowed configurationincluding the payload 800, payload attachment features 225, payloadchassis 210, mounting plate 430, PES to host connectors 440, launch lockassemblies 410, release mechanism 460, and deployment hinge assemblies500.

FIG. 9 is a section along the line 9-9 in FIG. 4 illustrating therelative positions of the major assemblies in the deployed configurationincluding the payload 800, payload assembly 200, payload attachmentfeatures 225, payload chassis 210, payload electrical box 240, mountingplate 430, PES to host connectors 440, launch lock assemblies 410,release mechanism 460, lock mounting plates 413, payload release plate610, snubber arms 650, snubbers 651 and deployment hinge assemblies 500with upper hinge plates 501, lower hinge plates 502, upper hinge pins503, mid hinge pins 504, lower hinge pins 505, deployments springs 508,and snubber shafts 540.

FIG. 10 is a view along arrows 10 in FIG. 9 and shows an underside viewof the payload assembly 200 with the payload chassis 210, retainingbolts 414, payload contacts 221, connector alignment features 220,payload electrical connectors 230, and payload electrical box 240.

FIG. 11 is a view along arrows 11 in FIG. 9 and shows an overhead viewof the mechanical mounting assembly 400 with the deployment hingeassemblies 500 and the payload release plate assembly 600 omitted forclarity. It shows the relative positions of the base plate 430, lockmounting plates 413, release plate contact 433, connector alignment pins434, payload to PEM connectors 435, circuit board 436 and PES to hostconnectors 440 a and 440 b. In this embodiment a system of electricalredundancy has been instituted for reliability resulting in two equaland separate electrical circuits for each task. In addition, the payloadharness 432 connecting the payload to the host 700 through the PES tohost connector 440 a is kept separate from the launch lock harness 431that connects the various mechanisms and sensors on the Payload Ejectionsystem 100 to the host spacecraft 700 through PES to host connector 440b.

FIG. 12 is an oblique view of one of the deployment hinge assemblies 500in the deployed configuration. It shows the relative positions of thebase plate 430, with the release plate contact 433 and lock mountingplate 413 and the lower hinge bracket 507. Mounted to the lower hingebracket 507 are the upper hinge plate 501, lower hinge plate 502, upperhinge pin 503, mid hinge pin 504, lower hinge pin 505, upper hingebracket 506 as part of the payload release plate 610, the deploymentsprings 508, upper hinge bearings 509, mid hinge bearings 510, lowerhinge bearings 511 and the snubber shaft 540. The payload release plate610 also includes the snubber arms 650 and snubbers 651.

FIG. 13 is a section view along line 9-9 of FIG. 4 of one of thedeployment hinge assemblies 500 in the deployed configuration. It showsthe relative positions of the base plate 430, and lock mounting plate413 and the lower hinge bracket 507. Mounted to the lower hinge bracket507 are the upper hinge plate 501, lower hinge plate 502, upper hingepin 503, mid hinge pin 504, lower hinge pin 505, upper hinge bracket 506as part of the payload release plate 610, the deployment springs 508,upper hinge bearings 509, mid hinge bearings 510 lower hinge bearings511 and the snubber shaft 540. The payload release plate 610 alsoincludes the snubber arms 650 and snubbers 651.

FIG. 14 is a section view along line 8-8 of FIG. 2 of one of thedeployment hinge assemblies 500 in the stowed configuration. It showsthe relative positions of the payload chassis 210, payload mountingfeatures 225, launch lock assembly 410, lock control harness 412, baseplate 430, and the lower hinge bracket 507. Mounted to the lower hingebracket 507 are the upper hinge plate 501, lower hinge plate 502, upperhinge pin 503, mid hinge pin 504, lower hinge pin 505, upper hingebracket 506 as part of the payload release plate 610, the deploymentsprings 508, upper hinge bearings 509, lower hinge bearings 511 and thesnubber shaft 540.

FIG. 15 is an oblique section view along line 15-15 of FIG. 2 of one ofthe deployment hinge assemblies 500 in the stowed configuration with thepayload assembly 200 omitted for clarity. It shows how the snubbers 651engage the snubber shaft 540 and the relative positions of the lockmounting plate 413, base plate 430, upper hinge plate 501, lower hingeplate 502, upper hinge pin 503, mid hinge pin 504, the deploymentsprings 508, mid hinge bearings 510, payload release plate 610 andsnubber arms 650.

FIG. 16 is a partial section view along line 16-16 of FIG. 2 of one ofthe launch lock assemblies 410 in the stowed configuration. It shows therelative positions of the payload chassis 210, launch lock assembly 410,lock release mechanism 411, lock control harness 412, lock mountingplate 413, retaining bolt 414, retraction spring 415 in compressedconfiguration, lock bolt housing 416, spring housing 417, load cell 418,load cell housing 419, load cell harness 420, base plate 430 andlaunchlock harness 431.

FIG. 17 is a partial section view along line 17-17 of FIG. 4 of one ofthe launch lock assemblies 410 in the deployed configuration. It showsthe relative positions of the lock release mechanism 411, lock controlharness 412, lock mounting plate 413, base plate 430 and launch lockharness 431.

FIG. 18 is a partial section view along line 18-18 of FIG. 4 of one ofthe launch lock assemblies 410 in the deployed configuration. It showsthe relative positions of the payload chassis 210, retaining bolt 414,retraction spring 415 in extended configuration, lock bolt housing 416,spring housing 417, load cell 418, load cell housing 419 and load cellharness 420.

FIG. 19 is a partial section view along line 19-19 of FIG. 2 of one ofthe deployment load paths in the stowed configuration. It shows therelative positions of the payload chassis 210, payload contact 221,payload release plate 610, release plate contact 433 and base plate 430.

FIG. 20 is a partial section view along line 8-8 of FIG. 2 showing therelease mechanism 460 in the stowed configuration. It shows the relativepositions of the release mechanism 460, release shaft 461, release nut462, washer 463, mounting plate 464, lock control harness 412, baseplate 430, payload to PEM connectors 435, circuit board 436, PEM harnesssockets 437, payload electrical connector 230, payload harness pins 231and payload electrical box 240.

FIG. 21 is a partial section view along line 9-9 of FIG. 4 of thepayload side of the release mechanism 460 in the deployed configuration.It shows the relative positions of the payload chassis 210, releaseshaft head 461 a, release nut 462, washer 463, payload electricalconnector 230, payload harness pins 231 and payload electrical box 240.

FIG. 22 is a partial section view along line 9-9 of FIG. 4 of the hostside of the release mechanism 460 in the deployed configuration. Itshows the relative positions of the base plate 430, release mechanism460, release shaft stub 461 b, mounting plate 464, lock control harness412, payload to PEM connectors 435, circuit board 436 and PEM harnesssockets 437.

FIG. 23 is an oblique view of the host-side electrical connectionsbetween the PEM 300 and the payload assembly 200 showing the payloadharness 432, connector alignment pins 434, payload to PEM connectors 435and circuit board 436.

FIG. 24 is an oblique view of the payload-side electrical connectionsbetween the PEM 300 and the payload assembly 200 showing the payloadelectrical connectors 230, connector alignment features 220 and thepayload electrical box 240.

FIGS. 25a, 25b, 25c, 25d, and 25e are a sequence of partial sectionviews along line 8-8 of FIG. 2 showing the deployment sequence as thedeployment hinge assemblies 500 accelerate the payload release plateassembly 600 and payload assembly 200 away from the mechanical mountingassembly 400 until the deployment hinge assemblies 500 reach the end oftheir travel and the payload assembly 200 separates from the payloadrelease plate assembly 600 and continues onward under its own inertia.

FIG. 25a shows the mechanism ready to be released. The launch locks 410are released, the deployment springs 508 are held in place by theretaining action of the release mechanism 460.

FIG. 25b shows that the mechanism has been activated. The releasemechanism 460 has activated and the deployment springs 508, no longerrestrained, are causing the upper hinge panel 501 and lower hinge panel502 to rotate. This causes the payload release plate assembly 600 toaccelerate away from the mechanical mounting assembly 400. At thispoint, release plate assembly 600 and the payload assembly 200 are heldtogether only by the acceleration of the mechanism.

FIG. 25c shows that the mechanism continues to accelerate the payloadrelease plate assembly 600 away from the mechanical mounting assembly400.

FIG. 25d shows the mechanism at the instant the deployment hingeassemblies 500 are at their full extension and have come to a completestop.

FIG. 25e shows the mechanism after the deployment hinge assemblies 500are at their full extension and have come to a complete stop and thepayload assembly 200 has separated from the deployment plate assembly600 and is free to move under its own inertia.

FIG. 26 is a block diagram showing host spacecraft 700 with acommunication antenna 701 for communicating with Earth 703 withspacecraft 700 having payload ejection system 100 mounted thereon incompartments such as unused battery bays 720 or an unused exteriorsurface area 710 of the spacecraft 700.

FIG. 27 is a block diagram of a non-limiting exemplary computer systemcoupled to the payload ejection system 100 which contains a centralprocessor 1210 interfaced with a memory storage device 1220,input/output devices and interfaces 1230, a power supply 1260, aninternal storage 1240 and a communications interface 1250.

DETAILED DESCRIPTION

Various embodiments and aspects of the disclosure will be described withreference to details discussed below. The following description anddrawings are illustrative of the disclosure and are not to be construedas limiting the disclosure. Numerous specific details are described toprovide a thorough understanding of various embodiments of the presentdisclosure. However, in certain instances, well-known or conventionaldetails are not described in order to provide a concise discussion ofembodiments of the present disclosure.

As used herein, the terms, “comprises” and “comprising” are to beconstrued as being inclusive and open ended, and not exclusive.Specifically, when used in the specification and claims, the terms,“comprises” and “comprising” and variations thereof mean the specifiedfeatures, steps or components are included. These terms are not to beinterpreted to exclude the presence of other features, steps orcomponents.

As used herein, the term “exemplary” means “serving as an example,instance, or illustration,” and should not be construed as preferred oradvantageous over other configurations disclosed herein.

As used herein, the terms “about” and “approximately”, when used inconjunction with ranges of dimensions of particles, compositions ofmixtures or other physical properties or characteristics, are meant tocover slight variations that may exist in the upper and lower limits ofthe ranges of dimensions so as to not exclude embodiments where onaverage most of the dimensions are satisfied but where statisticallydimensions may exist outside this region. It is not the intention toexclude embodiments such as these from the present disclosure.

As used herein, the term “operably connected” refers to a means ofcommunication between two devices. This can be either a wired ornon-wired communication.

As used herein, the term “tumble rate” is a toppling rotational rateabout any axis of a 3-axis orthogonal reference frame associated withthe centre of mass of the payload or payload assembly that isdetrimental to operation and/or recapture of the ejected payload.

Referring to FIG. 5, host spacecraft 700 often have surplus mass andvolume capacity on their exterior. As shown in FIG. 5, this can includeunused battery bays 720 or an unused exterior surface area 710 of thespacecraft 700 that could be used to host a payload ejection system 100,see FIG. 26. In one embodiment of this mechanism, it is proposed to usethese vacant spaces to house the payload ejection system 100 and itsattached payload 800. Other embodiments can include spacecraft that aredesigned specifically to carry and eject a plurality of payloads as partof their primary function as opposed to carrying payloads in addition totheir primary function.

As mentioned above, some existing methods of ejecting a payload from ahost spacecraft 700 in a microgravity environment such as orbit, applythe ejection force along a single vector and as such any displacement ofthe centre of mass of the payload from the vector of the ejection forceproduces a moment that is directly related to the distance between thecentre of mass and the vector and mass of the payload.

As illustrated in FIG. 6a , the ejection force 760 is applied along thedirection 780. The centre of mass 750 of the payload 800 is offset somedistance 790 from the direction 780. This combination of force at adistance produces a moment or couple 770 that causes the payload 800 torotate or tumble. The ejection mechanism itself cannot correct thiseffect and it requires that the payload either be manufactured with verystrict control of the location of the centre of mass, frequentlycompromising aspects of the payload, or the payload itself must expendresources to correct the tumble.

Similarly, other payload ejection methods rely upon the action of aplurality of springs 765 spread over a known area to provide theejection force as shown in FIG. 6b . In this case any distance betweenthe centre of mass 750 and the geometric centre of the group of springs795 means that springs closer to the centre of mass 750 exert theirforce against proportionally more of the payload 800 mass. This againcauses a moment or couple as the springs further from the centre of massextend faster and impart a rotation or tumble 770 to the payload 800.And, again, the ejection mechanism itself cannot correct this effectwith the same deleterious impacts on the payload.

There are several methods to mitigate the tumbling effect of an ejectedpayload. These include: ballasting the payload to collocate the centreof mass with the ejection force vector, and guiding the payload.Ballasting the payload is mass and volume expensive and requiresaccurate and specific knowledge of the mass properties of the payload.It must also be done uniquely for each payload decreasing operationalflexibility. Guiding the payload through the entire acceleration to theejection speed, as in the case of a PPOD, requires guides. Linear guidesare prone to jamming or binding as the payload approaches the end of theguides and the effective engagement of the guides is reduced to zero.

The present payload ejection system uses a plurality of deployment hingeassemblies 500 (FIG. 7) to eject a payload with a negligible amount ofinduced rotational rate or tumble even though the centre of mass of thepayload is, or may be significantly distant from the overall ejectionforce vector. A key to this mechanism is the use of two or more systemlinked hinges that produce parallel motion of one plane versus another.The payload ejection mechanism 300 uses at least two pairs of hingesconnected to two parallel planes and placed at an angle to each otherthus constraining the possible motion of the two planes relative to eachother to be parallel.

The present system can be readily scaled up to handle larger payloads byusing more than two deployment hinge assemblies 500. The payloadejection system 100 disclosed herein and illustrated has four (4)deployment hinge assemblies 500 but for larger payloads five, six, sevenand larger numbers of deployment hinge assemblies 500 may be used.Because the two hinges are at an angle they effectively describe aseries of parallel planes at each of the upper, mid and lower hinge axesconstraining the base plate 430 and the payload release plate 610 toremain parallel even in the presence of variations in the centre of massof the payload with respect to the geometric centre of the payloadrelease plate 610.

More specifically, FIG. 7 shows a simplified diagram of a deploymenthinge assembly 500 used in the payload ejection system 100. To minimisetorsional effects and reduce the required stiffness of the deploymenthinge assembly 500 the payload ejection system 100 uses a pair ofopposed linkages with each pair consisting of two upper hinge plates 501and two lower hinge plates 502 arranged orthogonally to each other. Themechanism in the figure shows each linkage hinge to bend outward aboutthe mid hinge pin 504, however to make the mechanism more compact thecurrent embodiment has one pair that bends inwards and another thatbends outwards. The direction of the hinge action has no bearing on theeffectiveness of the mechanism other than compactness and reduced mass.

For launch and any powered transit in the stowed configuration shown inFIG. 1 to the ejection site, the payload assembly 200 is secured to thebase plate 430 of the payload ejection mechanism 300 by one or morelaunch lock assemblies 410. The payload ejection mechanism 300 is in thestowed configuration and the deployment springs 508 (FIG. 9) are stowedin their maximum stored energy state.

When it is decided to eject the payload the launch lock assemblies 410are commanded to release and then the payload ejection mechanism 300 andthe deployment springs 508 are held by the release mechanism 460. At theappropriate time, the release mechanism 460 is commanded to release andwhen it does, the stored energy in the deployment springs 508 starts toforce the upper hinge panel 501 and lower hinge panel 502 to straightenup. The connector alignment pins 434 ensure that the payload electricalconnector 230 slides cleanly out of engagement with the payload to PEMconnector 434 before coming out of contact with the connector alignmentfeatures 220 themselves.

The actions of the pair of deployment hinge assemblies 500 drive thepayload release plate 610 away from the base plate 430 at a ratedetermined by the spring forces, the mechanism frictional drag and themass of the payload and with the payload release plate 610 remainingparallel to the base plate 430.

At the end of the travel of the deployment hinge assemblies 500 as shownin FIG. 12, the upper hinge plate 501 and the lower hinge plate 502 comeinto contact when the upper hard stop 530 contacts the lower hard stop531. The deployment spring 508 force then drops to zero and the payloadrelease plate 610 advances no further. The payload assembly 200 and theattached payload 800 are not physically attached to the release plateassembly 600, but payload assembly 200 is adjacent to payload releaseplate 610 in physical contact to form an interface between them but notin any way fixed to the payload release plate 610 so that payloadassembly 200 simply experiences the uni-axial ejection force created bythe deploying mechanism. At the point that the deployment hingeassemblies 500 reach their hard stops 530 and 531, the payload assembly200 becomes free of the payload release mechanism 300 which continues onthe ejection vector due to its own inertia, where its motion isperpendicular to the payload ejection plate 610 at time of release.

The mechanism will now be described in more detail with reference to thefigures.

At any time after the launch of the host spacecraft 700 and prior to thetime it is desired to eject the payload 800 and payload assembly 200 thecomputer control system 1200 either determines through internalprogramming or is commanded by a signal 702 from earth 703 to initiatethe payload ejection sequence. Prior to the issuance of the command toeject the payload being given by the computer control system 1200, thepayload ejection system 100 is in the stowed configuration as shown inFIGS. 1, 2, 8 and 11.

In this configuration, any power or data required by the payload ispassed from the host spacecraft 700 through the PES to host connectors440 a, the payload harness 432, the circuit board 436 to the PEM harnesssockets 437 held by the payload to PEM connectors 435. The power anddata then crosses to the payload assembly 200 via the payload harnesspins 231 held by the payload electrical connectors 230. A harnessconnects the payload harness pins 231 to the payload 800 and the payloadassembly 200. This harness is not shown because it is specific to eachcombination of payload 800 and payload assembly for each use of thepayload ejection system 100.

The commands to initiate payload 800 ejection are provided to orgenerated by the computer control system 1200 and passed to the payloadejection mechanism 300 via PES to host connectors 440 b and launch lockharness 431. The launch lock harness 431 provides a means to providepower and data connectivity to the launch lock assemblies 410 and therelease mechanism 460 and any sensors (not present in this embodiment)that may required for the operation and monitoring of the payloadejection mechanism 300.

Upon the command to operate the launch lock assemblies 410 and referringto FIGS. 16, 17 and 18 the signal and power from the launch lock harness431 passes to each the lock control harnesses 412. In this embodiment,the launch lock assemblies 410 are commercially available separable nutdevices. Upon command, the lock release mechanism 411 causes the nutwithin the lock release mechanism 411 to separate releasing theretaining bolt 414. The retraction spring 415 is also released and movesthe spring housing 417 and the retaining bolt 414 away from the baseplate 430 and up into the lock bolt housing 416, preventing theretaining bolt from causing the payload ejection system 100 from bindingor fouling.

Referring to FIGS. 20, 21 and 22, prior to initiation, the payloadejection system 100 is held together by that action of the releasemechanism 460 that prevents the deployment springs 508 from ejecting thepayload 800. At the appropriate time, as determined by programmingwithin the central computer system 1200 (see FIG. 27) or passed to thecentral computer system 1200 from earth 703 by signals 702 to the hostsatellite 700. see FIG. 26. The ejection command from the centralcomputer system 1200 is passed to the payload ejection mechanism 300 viathe PES 100 to host connectors 440 b (see FIG. 11) and the launch lockharness 432 which connects to the release mechanism 460.

In this embodiment, the release mechanism 460 is a commerciallyavailable frangible bolt device. Upon command the release mechanism 460causes the release shaft 461 to fracture in a precise manner leaving thebulk of the release shaft 461 b within the release mechanism 460attached to the mounting plate 464 and then to the base plate 430. Theremaining portion of the release shaft 461 a remains attached to therelease nut and attached to the payload assembly 200 during the ejectionsequence.

The deployment hinge assemblies 500 (described in detail below) push thepayload assembly 200 away from the payload ejection mechanism 300.Referring to FIGS. 20, 21, 22, 23 and 24 in order to provide a cleanrelease of the release mechanism 460, payload to PEM connectors 435 andcircuit board 436 are fixed to mounting plate 464 which is attached tobase plate 430 in such a way to permit limited movement in the plane ofthe base plate 430 and perpendicular to that plane. This movementremoves any stresses on the release mechanism 460 and electricalconnectors 230 and 435 that might prevent them from disengaging easily.To further guide the disengagement of the connectors 230 and 435 duringejection, the alignment of the payload electrical connectors 230 to thepayload to PEM connectors 435 is maintained by the connector alignmentpins 434 that are mounted releasably within the connector alignmentfeatures 220 that form a part of the payload electrical box 240. By thecombined action of close manufacturing tolerances and lubricatedsurfaces the connector alignment pins 434 slide easily within theconnector alignment features 220 and yet restrain unwanted movementbetween the payload electrical connectors 230 and the payload to PEMconnectors 435. Upon ejection, as the payload assembly 200 moves awayfrom the payload ejection mechanism, the payload harness pins 231 thatare part of the payload electrical connectors 230 slide out ofengagement of the PEM harness sockets 437 that are part of the PEMconnectors 435 while the connector alignment pins 434 are still engagedwithin the connector alignment features 220. After the payload harnesspins 231 have completely moved out of engagement with the PEM harnesssockets 437 the connector alignment pins 434 then move out of engagementwith the connector alignment features 220.

The deployment hinge assemblies 500 provide the force that enables theejection of the payload 800 and payload assembly 200. Referring to FIGS.12, 13, 14 and 15 the deployment hinge assemblies 500 work in thefollowing manner. As explained above, when the release mechanism 460(FIG. 22) is activated the payload release plate 610 is then free to beacted upon by the deployment hinge assemblies 500. Specifically, thedeployment springs 508 are configured to act upon the upper hinge plate501 and the lower hinge plate 502 in such a way as to force them fromthe collapsed or stowed configuration (FIG. 14) to the extended ordeployed configuration (FIG. 12). The configuration of the deploymenthinge assemblies 500, specifically the use of a system of two or morelinked hinge pairs produces parallel motion of one plane versus another.The deployment hinge assemblies 500 use at least two sets of hingesconnected to two parallel planes, the base plate 430 and the payloadrelease plate 610, and placed at an angle to each other thusconstraining the possible motion of the two planes to be parallel. Apreferred embodiment of the payload ejection system disclosed herein hasfour (4) deployment hinge assemblies 500 and any pair of adjacent,non-parallel deployment hinge assemblies 500 are sufficient to constrainthe motion of the payload release plate 610 to be parallel to the baseplate 430, however the use of additional deployment hinge assemblies 500reduces the torsional loads within the mechanism and reduces therequired stiffness of the deployment hinge assemblies 500 advantageouslyreducing the mechanism mass and increasing reliability.

As the deployment springs act upon the upper hinge plate 501 and thelower hinge plate 502 they rotate about the mid hinge pins 504 whichthen causes the upper hinge plate 501 to rotate around the upper hingepin 503 and the lower hinge plate 502 to rotate around the lower hingepin 505. The physical arrangement of one deployment hinge assembly 500in relationship to any adjacent deployment hinge assembly 500, ischaracterized by the two deployment hinge assemblies 500 being attachedto the payload deployment plate 610 and the base plate 430 such that

-   -   a) all of the upper hinge pins 503 are in one plane,    -   b) all of the mid hinge pins 504 are in a second plane and    -   c) all of the lower hinge pins 505 are in a third plane and that    -   d) the axes of all of these hinge pins (503, 504 and 505) form a        non-zero angle (in this case they are orthogonal) with those of        the adjacent deployment hinge assembly 500.

This means that the minimum two adjacent hinge assemblies effectivelydescribe a series of parallel planes at each of the upper, mid and lowerhinge axes, preventing the base plate 430 or the payload release plate610 from being pushed out of parallel as the deployment springs 508 actto extend the individual deployment hinge assemblies 500. Thisconstrained motion is what forces the payload release plate 610 to movein a manner parallel to the base plate 430 when (referring to FIG. 6a or6 b) even when the center of mass 750 of the payload 800 is asignificant distance 790 from the total ejection force vector 760 asapplied by the deployment hinge assemblies 500.

As the deployment hinge assemblies 500 reach their desired limit oftravel (refer to FIGS. 13 and 14) the upper hard stop 531, which is afeature on the upper hinge plate 501, comes into contact with the lowerhard stop 532 which is a feature on the lower hinge plate 502, and theextension of that deployment hinge assembly 500 comes to a stop. Due tothe arrangement of angularly arranged deployment hinge assemblies 500,each deployment hinge assembly 500 will come to its end of travel atsubstantially the same time therefore ending the ejection accelerationof the payload release plate 610 away from the base plate 430.

Referring to FIG. 19, in the stowed configuration there is no directloading between the payload contacts 221, the payload release plate 610and the release plate contact 433. Operation vibrations and loads maycause some contact between all three components and the release platecontact 433 is designed to restrict any excessive movement between thepayload release plate 610 yet remaining free of the payload releaseplate 610 in nominal conditions. Upon release mechanism 460 activation,as the deployment hinge assemblies 500 act to push the payload releaseplate 610 away from the base plate 430, the payload release plate 610now comes into firm contact with the payload contacts 221 at fourplaces. The acceleration of the payload assembly 200 provided by theactions of the deployment hinge assemblies 500 provides a force thatkeeps payload contacts 221 on the payload assembly 200 in controlledcontact with the payload release plate 610 during the ejection sequence.When the deployment hinge assemblies 500 have reached their full rangeof motion and no longer provide an acceleration, then the payloadcontacts 221 simply move away from the payload release plate 610 and thepayload assembly 200 and payload 800 are then independent of the hostsatellite 700.

It should be emphasised that the current payload ejection system 100does not require an additional latch device between the payload releaseplate 610 and the payload assembly 200 which would have to be timed torelease at or just before full extension of the PEM hinge assemblies500. This lack of a need for additional latches is enabled by thedeployment hinge assemblies 500 providing the uni-axial ejection forceand is predicated on the center of mass of the payload 800 and thepayload assembly 200 lying within the rectangle formed by the fourpayload contacts 221.

As described above, the connection of the PEM 300 to the payloadassembly 200 once the final release mechanism has been released isbetween the payload release plate 610 and the payload contacts 221. Thisconnection is a ‘push-contact’ only. This is chosen to ensure that oncethe ejection event has begun there is no risk that separation would notoccur. This then requires that the center of mass of the payloadassembly 200 is within the area contained by the payload contacts 221 onthe payload assembly 200 and payload release plate 610. This applies toall embodiments disclosed herein. Otherwise a tipping effect would occurregardless of the parallel motion provided by the ejection linkage. Itis possible to add a latch feature that would prevent this tipping ifthe center of mass was outside of this contact pattern, but the releaseof the latches would need to be timed so as not to interfere with thepayload assembly 200 at the moment of separation from the PEM 300.

FIG. 8 shows the payload ejection system 100 in its stowedconfiguration. The payload 800 can be virtually anything compatible withthe space environment. This includes, but is not limited to smallsatellites, satellite subcomponents, space system consumables such aspropellant or tools, components for the construction or maintenance ofspace systems, etc. The payload 800 can also be a unitary item or anaggregate of items fastened individually to the payload chassis 210using the payload attachment features 225. The payload attachmentfeatures 225 are simple threaded holes in this embodiment, however,depending upon the mission or the payload these features may also be aplurality of passive or active (motorized) attachment mechanisms each ofwhich facilitates the mechanical attachment of the payload(s) 800 plusproviding access to power, data and heat from the host 700 via cableharnesses that originate in the host 700 and pass to the payload via thepayload to host connectors 440 b, the payload harness 432, the circuitboard 436, the payload to PEM connectors 435, the payload electricalconnectors 230 and the mission specific harness(es) that would lead fromthe payload electrical connectors 230 to the payload 800. This is notshown as it would be unique to each payload.

In order to ensure that the mechanism does not bind during activation,several features have been incorporated in the payload ejection system100. Referring to FIG. 12, which shows the general arrangement of thedeployment hinge assemblies 500, the combination of deployment forceapplied by the deployment springs 508 coupled with maximum offsetdistance 790 the payload 800 centre of gravity 750 can be from thegeometric centre of the payload ejection mechanism 300 creates a momentor couple 770 that must be resisted by the deployment hinge assembly500. Through the choice of manufacturing tolerances and the stiffnessesof the hinge plate 501 and 502 and hinge bearing 509, 510 and 511, theinevitable flexing that happens within the mechanism can be accommodatedwhile minimising system mass and maximising the payload offset distance790, thereby maximising the system's utility.

Referring to FIG. 16 the launch lock assemblies 410 are configured tominimise the chances of the lock release mechanism 411 failing torelease the payload assembly 200 from the payload ejection mechanism300. In the stowed configuration the exact clamping force needed to holdthe payload assembly 200 to the payload ejection mechanism 300 isestablished during assembly by the use of a load cell 418 as one of theclamped components. The data from the load cell can be read duringassembly and the load cell harness 418 can be trimmed at that point ifcontinuous monitoring is not needed or the harness can be integratedinto the payload electrical connector 230 via the payload harness pins231.

When activated, the lock release mechanism 411 releases the retainingbolt 414 and the retraction spring 415 pulls the retaining bolt 414 backand up into the lock bolt housing 416, out of the way and minimising thechances of these bolts jamming the mechanism.

Referring to FIG. 20, to ensure the electrical connectors 230 and 435between the payload assembly 200 and the payload ejection mechanism 300separate cleanly the release mechanism 460 is rigidly fastened to themounting plate 464 but the mounting plate 464 has limited freedom ofmovement in the radial and axial directions. This permits the assemblyof parts rigidly held by the release mechanism 460 to accommodate themovement of the other parts of the payload ejection mechanism 300. Thisassembly of rigidly held parts includes the payload electrical box 240with the attached payload electrical connectors 230, payload harnesspins 231, payload to PEM connectors 435 with the attached PEM harnesssockets 437. To further ensure alignment of the connectors 230 and 435during separation the two connector alignment pins 434 slide within twoconnector alignment features 220 that are manufactured to tighttolerances to ensure binding does not occur.

Referring to FIG. 15, when the deployment hinge assemblies 500 arecollapsed in the stowed configuration there is some freedom of movementbetween the various elements of the mechanism. This freedom of movementcan cause deleterious effects during the phases of the mission prior tothe desired ejection of the payload assembly 200. This embodiment uses aseries of compliant snubbers 651 to restrict and damp out potentialelement movement prior to payload assembly 200 ejection. The snubbers651 are attached to the snubber arms 650 which are attached to thepayload release plate 610 and are configured such that when the payloadrelease mechanism is in the stowed configuration, there is a nominalinterference between the snubber shaft 540 and the snubbers 651. Thecompliant nature of the snubbers 651 results in a spring force beingapplied to the snubber shafts 540 which acts to restrict the motion ofthe mid hinge pins 504 and thereby restricts and secures the rest of thecomponents of the deployment hinge assemblies 500 preventing potentialdamage prior to the initiation of the command by the central computersystem 1200.

Referring to FIGS. 8 and 9, the launch lock assemblies 410 are theprimary structural connection between the payload assembly 200 and thepayload ejection mechanism 300 that withstands the forces generatedduring the hosting spacecraft's launch from earth and orbital manoeuvresup to the time that payload ejection is initiated in the desired orbit.

In an alternate embodiment, the release mechanism assembly 460 can bedesigned to be capable of bearing the launch loads entirely, such thatlaunch lock assemblies 410 would not be necessary. In this case, thestructure of the release mechanism assembly 460 would be configured toact as the primary structural load path and bear the loads generated inthe plane of the base plate 430 during spacecraft launch while the lockrelease mechanism 411 would provide the clamping load to react thelaunch loads perpendicular to the base plate 430.

There are several commercially available release mechanisms which may bechosen to be used for the launch release mechanism 411 or the releasemechanism 460. The choice of mechanisms depends on the requirements forthe mission. These mechanisms include frangible bolt systems, burnthrough mechanisms, separable nut systems and pyrotechnic systems, whichwill be well known by those skilled in the art. Key elements in thisembodiment are that the launch release mechanisms 411 are sized towithstand the launch structural loads and the release mechanism needs tobe sized only to hold back the deployment springs 508 prior to the finalcommand to eject the payload assembly 200.

An alternate embodiment would exchange the stored energy activation ofthe deployment springs 508 for a powered actuator(s) that drive thehinge plates 501 and 502 to deploy. Using a powered actuator can confera different acceleration profile to the payload assembly 200 which maybe advantageous in some situations or environments.

An alternate embodiment would add features to the payload assembly 200suitable to permit the ejected payload 800 and attached payload assembly200 to be grasped or captured by a device attached to a spacecraft forthe purpose that this captured payload may be attached to or used by thecapturing spacecraft. Payloads 800 where it might be desirable for themto be captured by a separate spacecraft would be payloads consistingspare parts, additional propellant, or mechanisms conferring additionalfeatures to the capturing spacecraft. It is in situations where thepayload assembly 200 will be captured by another spacecraft where thegreatly reduced tumble rates produced by the payload ejection system areespecially advantageous. Reduced payload assembly 200 tumble ratessignificantly reduce the difficulty of another spacecraft capturing theejected payload assembly 200.

Features that enhance or enable the capture of a payload assembly 200 byanother satellite include, but are not limited to, things such asgrapple features to enable the physical contact and capture between twospacecraft, visual or radar targets that enhance and enable manual orautomated visual, LIDAR and radar tracking by the capturing spacecraft,interface mechanisms that enable the captured payload assembly 200 to besecurely attached to the capturing spacecraft enabling the payload 800to be utilised.

Examples of some of the features usable for a spacecraft to capture thepayload assembly 200 are those used in the Orbital Express DemonstrationMission (Ogilvie, A., Autonomous Satellite Servicing Using the OrbitalExpress Demonstration Manipulator System, Proceedings of the 8thInternational Symposium on Artificial Intelligence, Robotics andAutomation in Space, iSAIRAS, Pasadena, 2008 and Ogilvie, A., AutonomousRobotic Operations for On-Orbit Satellite Servicing, Sensors and Systemsfor Space Applications, Proc. Of SPIE Vol 6958, 695809, 2008).

The present payload ejection system may be retrofitted onto any suitablesatellite to be used as a host spacecraft. The system may be underteleoperation by a remotely located operator, for example located onearth, in another spacecraft or in an orbiting space station. The systemmay also be autonomously controlled by a local Mission Manager with somelevels of supervised autonomy so that in addition to being under pureteleoperation there may be mixed teleoperation/supervised autonomy.

An alternate embodiment would add features that would permit the payloadejection mechanism 300 to be retracted after activation and change therelease mechanism 460 from a single use device such as the frangiblebolt devise to one that can be reset remotely. Retraction features mayinclude, but are not limited to, cables connected to a winch and motoror a piston and lever arrangement with appropriate hasps and latches.This would allow an additional device (not shown) to place additionalpayloads 800 and payload assemblies 200 upon the reset payload ejectionmechanism 300 so that these additional payloads 800 and payloadassemblies 200 may also be ejected. This is a useful embodiment in caseswhere multiple payloads are being launched with one payload ejectionmechanism having a first payload 800 coupled thereto but where additionpayloads 800 are stored on the host satellite and can be sequentiallyretrieved from their stored locations and ejected once the first payloadhas been ejected. An autoloader mounted on the host satellite may beprogrammed to fetch the additional payloads and mount them on thepayload deployment plate. The autoloader would be pre-programmed torelease the addition payloads from their storage berths. Optionally avision system may be positioned on the host satellite so the re-launchoperations may be controlled remotely by a human operator.

The specific embodiments described above have been shown by way ofexample, and it should be understood that these embodiments may besusceptible to various modifications and alternative forms. It should befurther understood that the claims are not intended to be limited to theparticular forms disclosed, but rather to cover all modifications,equivalents, and alternatives falling within the spirit and scope ofthis disclosure.

Therefore what is claimed is:
 1. A method for controllably ejecting apayload from a host spacecraft, comprising: attaching a payload to apayload assembly, releasably affixing the payload assembly to a payloadrelease plate of a payload ejection mechanism, said payload ejectionmechanism including a release mechanism; said payload ejection mechanismincluding at least two deployment assemblies coupled to said base platevia a first hinge coupling at one end thereof to the base and at anotherend thereof to said payload release plate via a second hinge coupling,said at least two deployment assemblies being extendable between astowed position and a fully extended position, said first and secondhinge coupling each having a hinge axis, said hinge axes being nonparallel to each other so that upon deployment said at least twodeployment assemblies are constrained to deploy such that a plane ofsaid payload assembly remains parallel to said base plate; wherein saidpayload ejection mechanism accommodates payloads having centers of massnot coincident with the geometric center of mass of said payload releaseplate in order to eject said payload assembly away from said hostspacecraft such that an effective force vector generated by said payloadejection mechanism acts through a center of mass of said payload at amoment of release of said payload assembly from said payload ejectionmechanism regardless of the location of said payload center of mass withrespect to the geometric center of said payload release plate; andejecting the payload assembly by activating the release mechanism todeploy said at least two deployment assemblies from their stowedpositions to their fully extended positions to force the payloadassembly away from the host satellite.
 2. The method according to claim1 wherein said at least two deployment assemblies are at least twodeployment hinge assemblies, said at least two deployment hingeassemblies each including two hinge plates hinged together along acommon mid-axis axis along the lengths of each of said hinge plates, andone hinge plate of each of said at least two deployment hinge assembliesbeing hinged to said base plate along said first hinge coupling anddefining a lower hinge axis and the other hinge plate being hinged tosaid payload release plate along said first hinge coupling and definingan upper hinge axis, and wherein each said common mid-axis, lower hingeaxis and upper hinge axis of a given deployment hinge assembly iscoplanar with each respective said common mid-axis, lower hinge axis andupper hinge axis of all other deployment hinge assemblies.
 3. The methodaccording to claim 1 wherein said payload ejection mechanism includes atleast four deployment hinge assemblies with a first of two pairs of saidat least four deployment hinge assemblies being in opposed relationship,and with a second of two pairs of said at least four deployment hingeassemblies also being in opposed relationship such said at least fourdeployment hinge assemblies form a rectangular shape.
 4. The methodaccording to claim 2 including at least one launch lock mechanism forrestraining and securing said payload ejection system and said payloadassembly until a commanded time of deployment at which time the launchlock mechanism is disengaged.
 5. The method according to claim 2 whereinsaid payload ejection mechanism includes an actuator to actuate saidrelease mechanism.
 6. The method according to claim 5 wherein saidactuator includes springs to actuate said payload ejection mechanism. 7.The method according to claim 5 wherein said actuator includes a motorto actuate said payload ejection mechanism.
 8. The method according toclaim 1 wherein said payload assembly is permanently attachable to apayload.
 9. The method according to claim 1 wherein said payloadassembly is releasably attachable to a payload.